Modern aircraft and spacecraft are fabricated from strong, lightweight composite materials. Such composite materials may include metals, metal alloys, polymers, glasses, carbon fibers, and the like. These materials may be formed as reinforced sandwich panels, in which two face sheets are held in a spaced-apart relationship by a reinforcing core. The core material is itself a rigid, lightweight material, such as a foamed or honeycomb structure. In use, aircraft and spacecraft may experience extremes of vibration, stress and temperature, and the structural members, including the sandwich panels and their fasteners, must be appropriate to such use. In the past, such panels have been fastened together by fasteners, such as bolts, or by the low-stress L-shaped joining pieces described in U.S. Pat. No. 5,212,003, issued May. 18, 1993 in the name of Homer. Another joining method is described in U.S. Pat. No. 5,324,146, issued Jun. 28, 1994 in the name of Parenti et al., in which the fasteners are assembled with adhesive in a manner which tends to take up tolerances which build up in the structure during assembly, which thereby tends to reduce the amount of assembly labor.
The requirements placed on such lightweight structures now dictate that, in some cases, the panels be tapered in thickness. In such a tapered-thickness panel, the face sheets are nonparallel, and the reinforcing core must also taper in thickness. It may be difficult to drill a clearance hole having a precise length through such a tapered-thickness panel to accommodate a fastener. Once such a clearance hole is made, the fastener may protrude, and the head may not lie flush with the panel.
Improved panel-to-panel and panel-to-equipment mounting arrangements are desired.